Intercooled cooling air with plural heat exchangers

ABSTRACT

A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passing the tapped air through a first heat exchanger and then to a cooling compressor. A second tap taps air from a location closer to the downstream most end than the location of the first tap, and the first and second taps mix together and are delivered into the high pressure turbine. The cooling compressor is positioned downstream of the first heat exchanger, and upstream of a location where air from the first and second taps mix together.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part of U.S. patent applicationSer. No. 14/695,534, filed Apr. 24, 2015.

BACKGROUND

This application relates to improvements in providing cooling air from acompressor section to a turbine section in a gas turbine engine.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as propulsion air. Further, the fan delivers air intoa compressor section where it is compressed. The compressed air passesinto a combustion section where it is mixed with fuel and ignited.Products of this combustion pass downstream over turbine rotors drivingthem to rotate.

It is known to provide cooling air from the compressor to the turbinesection to lower the operating temperatures in the turbine section andimprove overall engine operation. Typically, air from the downstreammost end of the compressor has been tapped, passed through a heatexchanger, which may sit in the bypass duct and then delivered into theturbine section. The air from the downstream most end of the compressorsection is at elevated temperatures.

SUMMARY

In a featured embodiment, a gas turbine engine comprises a maincompressor section having a high pressure compressor with a downstreammost end, and more upstream locations. A turbine section has a highpressure turbine. A first tap taps air from at least one of the moreupstream locations in the main compressor section, passing the tappedair through a first heat exchanger and then to a cooling compressor. Asecond tap taps air from a location closer to the downstream most endthan the location of the first tap, and the first and second taps mixtogether and are delivered into the high pressure turbine. The coolingcompressor is positioned downstream of the first heat exchanger, andupstream of a location where air from the first and second taps mixtogether.

In another embodiment according to the previous embodiment, there is asecond heat exchanger downstream of the cooling compressor, and airpasses through the second heat exchanger before being delivered into thehigh pressure turbine.

In another embodiment according to any of the previous embodiments, amain fan delivers bypass air into a bypass duct and into the maincompressor section. The first and second heat exchangers are positionedwithin the bypass duct to be cooled by bypass air.

In another embodiment according to any of the previous embodiments, airis tapped from a third tap downstream of the second heat exchanger, andis delivered to a component on the gas turbine engine to be cooled.

In another embodiment according to any of the previous embodiments, thelocation of the third tap is upstream of a location where the first andsecond taps mix together.

In another embodiment according to any of the previous embodiments, thecomponent to be cooled is a bearing compartment.

In another embodiment according to any of the previous embodiments, airtemperatures at the downstream most location of the high pressurecompressor are greater than or equal to 1350° F.

In another embodiment according to any of the previous embodiments, thesecond tap is at the downstream most end.

In another featured embodiment, a gas turbine engine comprises a maincompressor section having a high pressure compressor with a downstreammost end, and more upstream locations. A turbine section has a highpressure turbine. A first tap taps air from at least one of the moreupstream locations in the main compressor section, passing the tappedair through a first heat exchanger and then to a cooling compressor. Thecooling compressor compresses air downstream of the first heatexchanger. A second heat exchanger is downstream of the coolingcompressor. A second tap taps air from a location closer to thedownstream most end than the location of the first tap. Air from thefirst and second taps mix together and being delivered into the highpressure turbine. Air is tapped from a third tap downstream of thesecond heat exchanger, and delivered to a component on the gas turbineengine to be cooled.

In another embodiment according to the previous embodiment, a main fandelivers bypass air into a bypass duct and into the main compressorsection. The first and second heat exchangers are positioned within thebypass duct to be cooled by bypass air.

In another embodiment according to any of the previous embodiments, airtemperatures at the downstream most location of the high pressurecompressor are greater than or equal to 1350° F.

In another embodiment according to any of the previous embodiments, thesecond tap is at the downstream most end.

In another embodiment according to any of the previous embodiments, airfrom the first tap mixes with air from the second tap at a locationdownstream of the cooling compressor.

In another embodiment according to any of the previous embodiments, thelocation of the third tap is upstream of a location where the first andsecond taps mix together.

In another embodiment according to any of the previous embodiments, thecomponent to be cooled is a bearing compartment.

In another embodiment according to any of the previous embodiments, thebearing compartment supports a shaft adjacent to a combustor within thegas turbine engine.

In another embodiment according to any of the previous embodiments, airtemperatures at the downstream most location of the high pressurecompressor are greater than or equal to 1350° F.

In another embodiment according to any of the previous embodiments, thesecond tap is at the downstream most end.

In another embodiment according to any of the previous embodiments, airtemperatures at the downstream most location of the high pressurecompressor are greater than or equal to 1350° F.

These and other features may best be understood from the followingspecification and drawings, the following of which is a briefdescription.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows an embodiment of a gas turbine engine.

FIG. 2 shows a prior art engine.

FIG. 3 shows an embodiment of an engine.

FIG. 4 shows an embodiment of a system.

FIG. 5 shows a second embodiment of a system.

FIG. 6 shows another embodiment of a system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

Airflow through the core airflow path C is compressed by the lowpressure compressor 44 then by the high pressure compressor 52 mixedwith fuel and ignited in the combustor 56 to produce high speed exhaustgases that are then expanded through the high pressure turbine 54 andlow pressure turbine 46. The mid-turbine frame 58 includes vanes 60,which are in the core airflow path and function as an inlet guide vanefor the low pressure turbine 46. Utilizing the vane 60 of themid-turbine frame 58 as the inlet guide vane for low pressure turbine 46decreases the length of the low pressure turbine 46 without increasingthe axial length of the mid-turbine frame 58. Reducing or eliminatingthe number of vanes in the low pressure turbine 46 shortens the axiallength of the turbine section 28. Thus, the compactness of the gasturbine engine 20 is increased and a higher power density may beachieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 20fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 34. In another non-limiting example embodiment the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan blades 42 and the number of low pressure turbine rotors isbetween about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 34 in the lowpressure turbine 46 and the number of blades 42 in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

Gas turbine engines designs are seeking to increase overall efficiencyby generating higher overall pressure ratios. By achieving higheroverall pressure ratios, increased levels of performance and efficiencymay be achieved. However, challenges are raised in that the parts andcomponents associated with a high pressure turbine require additionalcooling air as the overall pressure ratio increases.

The example engine 20 utilizes air bleed 80 from an upstream portion ofthe compressor section 24 for use in cooling portions of the turbinesection 28. The air bleed is from a location upstream of the downstreamend 82 of the compressor section 24. The bleed air passes through a heatexchanger 84 to further cool the cooling air provided to the turbinesection 28. The air passing through heat exchanger 84 is cooled by thebypass air B. That is, heat exchanger 84 is positioned in the path ofbypass air B. As disclosed below, an auxiliary compressor may drive theflow.

A prior art approach to providing cooling air is illustrated in FIG. 2.An engine 90 incorporates a high pressure compressor 92 downstream ofthe low pressure compressor 94. As known, a fan 96 delivers air into abypass duct 98 and into the low pressure compressor 94. A downstreammost point 82 in the high pressure compressor 92 provides bleed air intoa heat exchanger 93. The heat exchanger is in the path of the bypass airin bypass duct 98, and is cooled. This high pressure high temperatureair from location 82 is delivered into a high pressure turbine 102.

The downstream most point 82 of the high pressure compressor 92 is knownas station 3. The temperature T3 and pressure P3 are both very high.

In future engines, T3 levels are expected to approach greater than orequal to 1350° F. Current heat exchanger technology is becoming alimiting factor as heat exchangers are made of materials, manufacturing,and design capability which have difficulty receiving such hightemperature levels.

FIG. 3 shows an engine 100 coming within the scope of this disclosure. Afan 104 may deliver air B into a bypass duct 105 and into a low pressurecompressor 106. High pressure compressor 108 is positioned downstream ofthe low pressure compressor 106. A bleed 110 taps air from a locationupstream of the downstream most end 82 of the high pressure compressor108. This air is at temperatures and pressures which are much lower thanT3/P3. The air tapped at 110 passes through a heat exchanger 112 whichsits in the bypass duct 105 receiving air B. Further, the air from theheat exchanger 112 passes through a compressor 114, and then into aconduit 115 leading to a high turbine 117. This structure is all shownschematically.

Since the air tapped at point 110 is at much lower pressures andtemperatures than the FIG. 2 prior art, currently available heatexchanger materials and technology may be utilized. This air is thencompressed by compressor 114 to a higher pressure level such that itwill be able to flow into the high pressure turbine 117.

An auxiliary fan 116 is illustrated upstream of the heat exchanger 112.The main fan 104 may not provide sufficient pressure to drive sufficientair across the heat exchanger 112. The auxiliary fan will ensure thereis adequate air flow in the circumferential location of the heatexchanger 112.

In one embodiment, the auxiliary fan may be variable speed, with thespeed of the fan varied to control the temperature of the air downstreamof the heat exchanger 112. As an example, the speed of the auxiliary fanmay be varied based upon the operating power of the overall engine.

Details of the basic system are disclosed in co-pending U.S. patentapplication Ser. No. 14/695,578, entitled “Intercooled Cooling Air,” andfiled on Apr. 24, 2015.

FIG. 4 shows an engine 121 that provides a variation to the basic systemdisclosed above. A bypass duct 120 again receives a heat exchanger 122,which communicates with the compressor 124. Air is tapped from the lowpressure compressor 128 at a location 126 and passes through the heatexchanger 122. A line 130 downstream of the compressor 124 communicateswith a mixer 132. Air is tapped at 140 from the most downstream locationin a high pressure compressor 141. While the most downstream location isdisclosed, it is possible to use an alternative location that is closerto the most downstream end than the location 126. The air is mixed withthe cooler air from line 130 in the mixer 132 and then passes into aline 142 to be delivered to the high pressure turbine 144.

With this arrangement, the amount of cooled air tapped from the lowpressure compressor 128 is reduced, and therefore the size of thecompressor 124 can be reduced.

FIG. 5 shows yet another embodiment wherein the mixer 132 is replaced bya valve 150 which is operable to vary the volume of air from tap 140 andthe line 130 being delivered into the line 142. Thus, at certain periodsof operation, a control 152 for the valve 150 may mix more or less ofthe hot air from tap 140 relative to the cooler air. As an example, atlower power operation, such as cruise, more of the hot air may be mixedwith the cool air. On the other hand, at high power operation, such astake-off, the percentage of cooler air from line 130 will be greaterrelative to the hot air from tap 140.

Details of the systems of FIGS. 4 and 5 are disclosed in co-pending U.S.patent application Ser. No. 14/695,534, filed on Apr. 24, 2015, andentitled “Intercooled Cooling Air Tapped From Plural Locations.”

FIG. 6 shows a system 200 that has two features. As shown, air 202 istapped, such as from the location 126. Air 202 passes through a firstheat exchanger 204, which sits in a cooling duct such as bypass duct120. That air passes to the auxiliary compressor 206. In this system, aplural, or “topping,” heat exchanger 208 is downstream of the compressor206. That air passes downstream and is mixed with hot air 209 such asfrom location 140 at a mixer 210. This system can also be incorporatedinto a system having the features of FIG. 5. Downstream of the mixer 210the air is delivered to the high pressure turbine at line 212.

The topping heat exchanger 208 also sits in the bypass duct 120, and theamount of cooling provided to the air will be increased. Thus, theamount of air moving through the auxiliary compressor 206, and all theplumbing connections can be made much smaller. Since the air downstreamof heat exchanger 208 is cooler, a higher percentage of air from line209 can be mixed compared to the air from tap 202.

Another optional feature is shown at 214 in FIG. 6. 214 is a tap fortapping air 216 for cooling purposes. Air downstream of the second ortopping heat exchanger 208 is cooled to lower levels than in the FIGS. 4and 5 embodiments, and thus the air is useful as cooling air such as fora bearing compartment 218. Bearing compartment 218 may be the #4 bearingcompartment, such as associated with supporting a shaft adjacent thecombustor.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A gas turbine engine comprising; a maincompressor section having a high pressure compressor with a downstreammost end, and more upstream locations; a turbine section having a highpressure turbine; a first tap tapping air from at least one of said moreupstream locations in said main compressor section, passing said tappedair through a first heat exchanger and then to a cooling compressor; asecond tap tapping air from a location in said main compressor sectioncloser to said downstream most end than the location of said first tap,and said first and second taps mixing together and being delivered intosaid high pressure turbine; said cooling compressor being positioneddownstream of said first heat exchanger, and upstream of a locationwhere air from said first and second taps mix together; wherein there isa second heat exchanger downstream of said cooling compressor, and airpasses through said second heat exchanger before being delivered to saidhigh pressure turbine; wherein air is tapped from a third tap downstreamof said second heat exchanger, and delivered to a component on the gasturbine engine to be cooled; and wherein the location of said third tapis upstream of a location where said first and second taps mix together.2. The gas turbine engine as set forth in claim 1, wherein saidcomponent to be cooled is a bearing compartment.
 3. A gas turbine enginecomprising; a main compressor section having a high pressure compressorwith a downstream most end, and more upstream locations; a turbinesection having a high pressure turbine; a first tap tapping air from atleast one of said more upstream locations in said main compressorsection, passing said tapped air through a first heat exchanger and thento a cooling compressor, said cooling compressor compressing airdownstream of said first heat exchanger; a second heat exchangerdownstream of said cooling compressor; a second tap tapping air from alocation in said main compressor section closer to said downstream mostend than the location of said first tap, wherein air from said first andsecond taps mixing together and being delivered into said high pressureturbine; air being tapped from a third tap downstream of said secondheat exchanger, and delivered to a component on the gas turbine engineto be cooled; and wherein the location of said third tap is upstream ofa location where said first and second taps mix together.
 4. The gasturbine engine as set forth in claim 3, wherein a main fan deliversbypass air into a bypass duct and into said main compressor section andsaid first and second heat exchangers are positioned within said bypassduct to be cooled by bypass air.
 5. The gas turbine engine as set forthin claim 4, wherein air temperatures at said downstream most location ofsaid high pressure compressor are greater than or equal to 1350° F. 6.The gas turbine engine as set forth in claim 4, wherein the second tapis at said downstream most end.
 7. The gas turbine engine as set forthin claim 4, wherein air from said first tap mixes with air from saidsecond tap at a location downstream of said cooling compressor.
 8. Thegas turbine engine as set forth in claim 3, wherein said component to becooled is a bearing compartment.
 9. The gas turbine engine as set forthin claim 8, wherein said bearing compartment supports a shaft adjacentto a combustor within said gas turbine engine.
 10. The gas turbineengine as set forth in claim 9, wherein air temperatures at saiddownstream most location of said high pressure compressor are greaterthan or equal to 1350° F.
 11. The gas turbine engine as set forth inclaim 3, wherein the second tap is at said downstream most end.
 12. Thegas turbine engine as set forth in claim 3, wherein air temperatures atsaid downstream most location of said high pressure compressor aregreater than or equal to 1350° F.